Rocket engine member and a method for manufacturing a rocket engine member

ABSTRACT

Method and arrangement for providing liquid fuel rocket engine member ( 10 ). The member forms a body of revolution having an axis of revolution and a cross section that varies in diameter along said axis. The wall structure includes a plurality of cooling channels ( 11 ). The outside of the wall structure includes a continuous sheet metal wall ( 14 ). The cooling channels ( 11 ) are longitudinally attached to the inside of the sheet metal wall. The method for manufacturing the rocket engine member ( 10 ) includes the steps of forming a sheet metal wall ( 14 ) having a wall section corresponding to the desired nozzle section, providing a plurality of channel members ( 15 ), and attaching the channel members ( 15 ) to the inside of the sheet metal wall ( 14 ).

CROSS REFERENCE TO RELATED APPLICATIONS

[0001] This application is a continuation patent application ofInternational Application No. PCT/SE02/00026 filed 9 Jan. 2002 which waspublished in English pursuant to Article 21(2) of the Patent CooperationTreaty, and which claims priority to Swedish Application No. 0100079-3filed 11 Jan. 2001 and to U.S. Provisional Application No. 60/261,044filed 11 Jan. 2001. Said applications are expressly incorporated hereinby reference in their entireties.

BACKGROUND OF INVENTION

[0002] 1. Technical Field

[0003] The present invention relates to a rocket engine.

[0004] 2. Background of the Invention

[0005] During operation, a rocket combustion chamber or a rocket enginenozzle is subjected to very high stresses, for example in the form of avery high temperature on its inside (on the order of magnitude of 980°F.) and a very low temperature on its outside (on the order of magnitudeof −370° F.). As a result of this high thermal load, stringentrequirements are placed upon the choice of material, design andmanufacture of the outlet nozzle. At a minimum, the need for effectivecooling of the combustion chamber or the outlet nozzle must beconsidered.

[0006] It is a problem to construct cooled wall structures that arecapable of containing and accelerating the hot exhaust gas, and also bereliable through a large number of operational cycles. Known designs donot have a sufficiently long service life required to withstand a largenumber of operational cycles. These known systems generate large thermalstresses, including large pressure drops, or present difficulties whenneeding repair.

[0007] When applying the expander engine cycle, there is a secondaryproblem. The expander engine cycle uses the cooling medium to drive theturbines in the fuel and oxidator turbo pumps; that is, energy from theexpansion of the heated cooling medium is used for driving the turbines.The efficiency of the rocket engine is a function of the combustionpressure. To reach high pressure experience in the expander cycle,efficient heat transfer from the exhaust gas to the cooling medium isrequired. Increase in the heat load in the combustion chamber due tosurface roughness or fins may impair the service life of the enginesince the intensity of the heat load is very high in the combustionchamber. Still further, a longer combustion chamber increases the lengthof the engine and the rocket. A commensurate increase in the size of thenozzle gives rise to larger nozzles and longer rocket structures, eachof which increases the weight of the vehicle.

[0008] There are several different known methods for manufacturing arocket nozzle with cooling channels. According to one of these methods,the nozzle has a brazed tube wall. The tubes have a varyingcross-sectional width to provide the contour of the nozzle whenassembled. The variation in cross section is given by variation of thecircumference and by variation of the form of the cross section. Thebrazed joints restrict the deformation of tubes in the thermal expansionand pressure cycle. The stresses in the tubes are increased in the arcof the joints. The joints themselves are weak points that may break andare difficult to repair. The brazed tube wall provides a larger “wet”contact surface for the rocket flame than a sandwich wall or a constanttube section wall. However, even larger wet surfaces are desirable.

[0009] According to another known method, a sandwich wall is made bymilling channels in sheet metal and joining a thinner sheet metal toseal the channels. The inner and outer walls are continuous shells. Inthe thermal cycle, the walls are in compressive and tension strain. Thistype of wall structure is not well suited to sustain the tension loadsnormal in the service life of a rocket nozzle. The sandwich wallfeatures no increase in surface area to enhance heat transfer.

[0010] According to still another known method, the nozzle wall ismanufactured with constant section tubes. The tubes are helically woundand welded together to form the nozzle contour. The increase in surfacearea is small. The tubes have an angle relative to exhaust gas flowingthrough the nozzle. This helps to increase the heat transfer. However,at the same time the exhaust flow is rotated and a reactive rollmomentum influences the flight of the rocket. The constant section tubesresult in a large pressure drop that is not favorable for convectivelycooled engines. The large pressure drop is negative for the expandercycle type engine.

SUMMARY OF INVENTION

[0011] An objective of the present invention is to provide an improvedmethod for manufacturing a cooled rocket engine member, as well as theresulting member itself.

[0012] The objective is exemplarily achieved by means of the rocketengine member having an outside wall structure that includes acontinuous sheet metal wall, and that included cooling channels areformed by elongated elements that are longitudinally attached to theinside of the sheet metal wall.

[0013] According to the teachings of the invention, a rocket enginemember may be manufactured which presents high pressure capacity, and alow pressure drop, a long cyclic life, as well as advantageous arearatio. The manufacturing lead time and cost may also be optimized.

BRIEF DESCRIPTION OF DRAWINGS

[0014] The invention will be further described in the following, in anon-limiting way with reference to the accompanying drawings in which:

[0015]FIG. 1 is a schematic perspective view showing a nozzle configuredaccording to the teaching of the present invention;

[0016]FIG. 2 is a partial sectional view taken along the line 2-2 inFIG. 1, showing three cooling channels at the inlet end of the nozzle;

[0017]FIG. 3 is a similar view as FIG. 2, but showing two of the coolingchannels taken along the line 3-3 at the outlet end of the nozzle;

[0018]FIG. 4 is a partial sectional view from a position correspondingto that shown in FIG. 2, but illustrating an alternative embodiment ofthree cooling channels at the inlet end of a nozzle; according to asecond embodiment of the invention;

[0019]FIG. 5 is a similar view as FIG. 3, but shows cooling channels atthe outlet end of the nozzle of the embodiment of FIG. 4;

[0020]FIG. 6 is a similar view as FIG. 4, but showing another variationof the invention;

[0021]FIG. 7 is a partial sectional view showing cooling channels at theinlet end of a nozzle according to another embodiment of the invention;

[0022]FIG. 8 is a related view of the embodiment of FIG. 7 showingcooling channels at the outlet end of the nozzle; and

[0023]FIG. 9 is a cross-sectional view of yet another alternativeembodiment of the invention.

DETAILED DESCRIPTION

[0024]FIG. 1 shows a diagrammatic and somewhat simplified perspectiveview of an outlet nozzle 10 that is produced according to the teachingsof the present invention(s). The nozzle is intended for use in rocketengines of the type using liquid fuel, for example liquid hydrogen. Theworking of such a rocket engine is conventional and therefore notdescribed in detail. The nozzle 10 is cooled with the aid of a coolingmedium that is preferably also used as fuel in the particular rocketengine. The invention is, however, not limited to rocket engine membersof this type, but can also be used in combustion chambers and in thosecases in which the cooling medium is dumped after it has been used forcooling.

[0025] The outlet nozzle is manufactured with an outer shape that issubstantially bell-shaped. Thus, the nozzle 10 forms a body ofrevolution having an axis of revolution and a cross section that variesin diameter along said axis.

[0026] The nozzle wall is a structure comprising (included, but notlimited to) a plurality of mutually adjacent, tubular cooling channels11 extending substantially in parallel to the longitudinal axis of thenozzle from an inlet end 12 outlet to/on end 13 of the nozzle. Theoutside of the structure includes a continuous sheet metal wall 14. Thecooling channels 11 are formed by elongated elements in the form oftubes 15, that are curved in the longitudinal direction to conform tothe nozzle contour and are oriented axially along the nozzle wall. Inthis position, they are jointed to the metal wall by welding. The weldsare preferably made by laser welding from the outside. This assemblyforms a leak tight nozzle with all joints at the cool side of thestructure. Further, there is no joint, or weld, attaching two adjacenttubes to each other.

[0027] The cooling channels 11 in the embodiment illustrated in FIGS. 2and 3 are formed by circular tubes 15 having a varying cross section.The tubes 15 may be seamless and have a smaller cross section at theinlet end 12 of the nozzle than at the opposite outlet end 13. Eachelongated element 15 preferably delimits only one cooling channel.

[0028] The cooling tubes 15 are mounted without gaps therebetween. Atthe inlet end 12 of the nozzle, the thickness of the tube material isthin to minimize the maximum temperature and to allow the tubes to beflexible to deformation of the cross section. At the outlet end 13 ofthe nozzle, the tubes have a larger cross section, as well as a thickertube material. This variation in material thickness allows the tubes toadapt to increased pressure inside the tubes when the cooling mediumcontained therein expands. At the inlet, the tubes may be formed in anoval shape to increase the number of tubes.

[0029] The variations in tube cross section and tube material thicknessare gradual in the longitudinal direction of the nozzle.

[0030]FIGS. 4 and 5 show a second embodiment of the invention that isadapted for enhanced heat pickup. The cooling tubes 15 are manufacturedwith a constant material thickness and a gradually increasing diameter.The tubes have a smaller cross section at the inlet end 12 of the nozzlethan at the opposite end. The inlet ends of the tubes have machinedfaces to allow a small pitch at this end of the nozzle to enable largearea ratios. The cooling tubes are mounted without mutual gaps at theinlet end of the nozzle where the flame pressure and heat load isgreatest.

[0031] At the outlet end 13 of the nozzle, the tubes 15 are separated inthe tangential direction (provided with mutual gaps therebetween). Acavity 16 is formed between each pair of tubes 15 and the sheet metalwall 14. The gap between the tubes allows the hot rocket flame to accessthe cavity and thus more tube surface for enhanced heat pick up. Also,by allowing a gap between each pair of adjacent tubes, the tube may beconical and yet be acceptable to fit a bell shaped nozzle. The variationin width of the cavity 16 between two adjacent tubes is gradual in thelongitudinal direction of the nozzle.

[0032] With the nozzle design described above, the amount of heattransferred to the coolant in the nozzle can be increased by a factor ofas much as 1.5 compared to conventional designs.

[0033] In cases where the heat load is high at the exit of the nozzle,the embodiment of the invention shown in FIG. 6 affects protection ofthe sheet metal wall 14 from the heat load. Thus, the cooling cavity maybe filled with a thermally insulating material 17 to prevent the gasfrom contacting the load carrying outer shell so that the shell materialtemperature is limited. Alternatively, the walls may be coated with athermally conductive material 17 for increased heat transfer to thecooling tubes. In a case where conductive material such as coppercompletely fills the cavity, it is possible to reach very high pressuresand high area ratios. The process to apply the conductive material canbe exemplarily be brazing or laser sintering.

[0034]FIGS. 7 and 8 show another embodiment of the invention whereU-formed profiles 18 are used instead of the above described circulartubes 15. The profiles have a varying cross section and a varyingmaterial thickness. The profiles are manufactured by press forming sheetmetal strips. The variation in thickness is adapted to the length of thenozzle. Thus, the material thickness may increase when the coolingchannel cross section is increased so that the thickness is small at theinlet end of the nozzle where the heat load is high. It is preferable tomodify the metal strip thickness before folding. The structure in FIGS.7 and 8 has been combined with the thermally insulating/conductivematerial 17.

[0035] It is possible to build the structures described above from thecommon materials for rocket engine nozzle tubes such as stainless steeland nickel based alloys. Also copper and aluminum are suitablematerials.

[0036] One of the important advantages of wall structures configuredaccording to the teachings of the present invention is that it offers alarge cooling surface for increased heat absorption. The variations incross section and tube material thickness allows for high internalpressure in the cooling channels 11. The increased wet surface, that is,the surface toward the exhaust gasses in the nozzle structure providedby the several embodiments of the invention(s) cools the boundary layermore than by a conventionally designed nozzle. The boundary layerleaving the disclosed rocket nozzle(s) will be cooler. The coolerboundary layer serves as cooling film for an eventual radiation coolednozzle extension that may be used as a low cost solution when the heatload is limited. The nozzle extension could be less costly since theheat load is limited.

[0037] The rotational symmetric outer surface of the nozzle structure(s)configured according to the teachings of the invention(s) also providesstiffness, and if necessary, allows for attachment of stiffeners in aneasy way. The single joint to the sheet metal wall isolates jackets andallows the tubes to be flexible to thermal distortion while imposing aminimum of stress concentration. The cross section of the coolingchannels may be close to circular. This means that the temperaturedifferences and associated stresses are lower than compared to sandwichwalls where the flame is not in contact with the outer wall. The gap 16between the tubes eliminates the restriction on cooling channeldimensions to form the nozzle contour. The cooling channels or tubescould be made with liner variation, which offers advantages inmanufacturing.

[0038] In FIG. 9, a further embodiment of the invention is shown in apartly cut away, cross-sectional view. A plurality of elongated elements21 are arranged next to each other. Each of the elongated elements 21has a plate-like portion, or base portion 23, and a plurality of flanges24, or ribs, which project from and extend along said base portion 23.The ribs 24 are elongated, arranged at a distance from each other andsubstantially in parallel to each other. Further, the elongated elementsare attached to a continuous sheet metal wall 14. Cooling channels 22are formed between two adjacent ribs 24 and the sheet metal wall 14.Further, two adjacent elongated elements are connected to each other bya further weld 25.

[0039] It should be appreciated that the invention is not limited to theabove-described embodiments, but modifications are possible while stillremaining within the scope of the presented claims.

1. A liquid fuel rocket engine member (10) comprising: a body having anaxis of revolution and a cross section that varies in diameter alongsaid axis, said body further comprising a wall structure having aplurality of cooling channels (11,19,20); and an outside of the wallstructure comprising a continuous sheet metal wall (14) and the coolingchannels (11) being at least partly delimited by elongated elements(15,18,21) that are longitudinally attached to the inside of the sheetmetal wall, the elongated elements (15,18) being mounted with mutualcontact at the inlet end (12) of the member and with mutual distances atthe outlet end (13) of the member.
 2. The liquid fuel rocket enginemember as recited in claim 1, further comprising: a cross sectional areaof each cooling channel being larger in a downstream end (13) of thechannel than in an upstream end (12) of the channel.
 3. The liquid fuelrocket engine member as recited in claim 1, further comprising: amaterial thickness of the cooling channel wall being larger in adownstream end (13) of the channel than in an upstream end (12).
 4. Theliquid fuel rocket engine member as recited in claim 1, furthercomprising: a width of each of said cooling channels, in thecircumferential direction of said engine member, being larger in adownstream end (13) of the channel than in an upstream end (12) of thechannel.
 5. The liquid fuel rocket engine member as recited in claim 1,further comprising: the cooling channels (11) having a substantiallysimilar cross section shape in a downstream end (13) of the channel asin an upstream end (12) of the channel.
 6. The liquid fuel rocket enginemember as recited in claim 1, further comprising: each cooling channel(11) being formed by a sheet metal profile (18).
 7. The liquid fuelrocket engine member as recited in claim 1, further comprising: thecooling channels being formed by seamless tubes (15).
 8. The liquid fuelrocket engine member as recited in claim 1, further comprising: adistance between two adjacent elongated elements (15,18) at the outletend (13) of the member (10) being filled with an insulating material(17).
 9. The liquid fuel rocket engine member as recited in claim 1,further comprising: the distance between two adjacent elongated elements(15,18) at the outlet end (13) of the member (10) being filled with athermally conductive material (17).
 10. The liquid fuel rocket enginemember as recited in claim 1, further comprising: at least one of saidelongated elements (21) defining a plurality of cooling channels (22).11. The liquid fuel rocket engine member as recited in claim 1, furthercomprising: the elongated element (21) being formed by a plate-shapedbase portion (23) and a plurality of flanges (24) projecting from saidbase portion, said cooling channels being formed between the baseportion, adjacent flanges and said sheet metal wall.
 12. A method formanufacturing a liquid fuel rocket engine member (10) comprising:providing a body having an axis of revolution and a cross section thatvaries in diameter along said axis, and the body having a wall structurecomprising a plurality of cooling channels (11); attaching a pluralityof elongated elements to a curved sheet metal wall arranged to form theengine member, wherein the cooling channels are formed by at least saidelongated elements and wherein the elongated elements (15,18) aremounted with mutual contact at the inlet end (12) of the member and withmutual distances at the outlet end (13) of the member.
 13. The method asrecited in claim 12; further comprising: forming a sheet metal wall (14)having a wall section corresponding to the desired member section. 14.The method as recited in claim 12; further comprising: defining thecooling channels by said sheet metal wall (14).
 15. The method asrecited in claim 12; further comprising: attaching the cooling channels(11) to the sheet metal wall (14) by welding from the outside of thewall.